1. Field of the Invention
The present invention relates generally to spacecraft and more particularly to spacecraft configurations.
2. Description of the Related Art
FIGS. 1 and 2 illustrate conventional spacecraft configurations which are described in various spacecraft references (e.g., see Morgan, Walter L., et al., Communications Satellite Handbook, John Wiley and Sons, New York, 1989, pp. 547-550 and Moral, G., et al., Satellite Communications Systems, John Wiley and Sons, New York, 1996, pp. 511-520). In particular, FIG. 1 illustrates an exemplary spin-stabilized spacecraft 20 and FIG. 2 illustrates an exemplary body-stabilized spacecraft 40.
In spin-stabilized spacecraft, all or a significant portion of the spacecraft spins about a spin axis. When an entire spacecraft spins, it is generally referred to as a "spinner". In contrast, spacecraft which have a nonspinning shelf portion are referred to as "dual spinners" or "gyrostats". For example, FIG. 1 shows a gyrostat spacecraft 20 which has a spinning drum 22 and a despun shelf 24. The drum spins about a spin axis 25 and is typically cylindrical with its outer surface carrying a solar cell array 26. In addition to this array, the drum 22 usually contains batteries, fuel tanks and thrusters.
The spacecraft's payload is generally mounted on the despun shelf 24. In FIG. 1, this payload is symbolized by broken lines 28 and a payload portion is exemplified by an antenna 30 which has a feed assembly 31 that transmits and receives signals 32 that are reflected off of a reflector 33.
The reflector 33 is carried at the end of a boom 34.
In order for spacecraft to meet their design objectives, their attitudes must typically remain within predetermined attitude envelopes. However, spacecraft are typically subjected to disturbance torques from a variety of external sources (e.g., solar pressures and gravity gradients). Because these disturbance torques will eventually rotate a spacecraft out of its attitude envelope, spacecraft attitude is generally controlled with control systems that include onboard torque generators (e.g., thrusters). The efficiency and lifetime of these control systems are enhanced if the spacecraft has a degree of natural stabilization.
Spin-stabilized spacecraft obtain their attitude stabilization via the gyroscopic principal. A disturbance torque T that is applied to a nonspinning spacecraft along an axis about which the spacecraft has a moment of inertia I will urge the spacecraft to rotate about that axis with a constant angular acceleration T/I. This angular acceleration will quickly rotate the spacecraft to the limit of its attitude envelope and, therefore, quickly require corrective control action.
In contrast, if the gyrostat 20 of FIG. 1 has a angular momentum H about its spin axis 25 and is subjected to the same disturbance torque T about an axis that is orthogonal to the spin axis, the spacecraft will rotate about a third orthogonal axis with an angular velocity T/H. Because an acceleration has been converted to a velocity and because the angular momentum H can be made quite large, the time before a corrective action is required has been greatly extended.
The body-stabilized spacecraft 40 of FIG. 2 has a pair of solar panels 41 and 42 which extend away from a nonrotating body 44. One side of each of the solar panels is covered with a solar cell array 46. The satellite's payload is carried within or on the body 44. In the case of a communications spacecraft, the payload typically includes antennas 48.
The attitude of a body-stabilized spacecraft which moves in an orbital plane about a celestial body is often referenced to an orthogonal coordinate system having pitch, yaw and roll axes. The pitch axis is orthogonal to the orbital plane, the yaw axis points to the celestial body and the roll axis is aligned with the spacecraft's velocity vector. In a geostationary orbit, the solar wings 41 and 42 are typically oriented along the pitch axis.
The body-stabilized spacecraft 40 lacks the attitude stabilization that the spin-stabilized spacecraft 20 of FIG. 1 gains from its spinning body. This stabilization is replaced with an attitude control system that typically includes flywheels whose rotation provides an on-board angular momentum.
FIGS. 3A-3C show schematized views of the spacecraft 40 of FIG. 2 with FIGS. 2 and 3A-3C having like elements indicated by like reference numbers. These figures illustrate exemplary stabilization elements for body-stabilized spacecraft.
In FIG. 3A, an attitude control system includes a single spinning flywheel 50 which is typically referred to as a momentum wheel. The angular momentum H of the momentum wheel 50 provides a measure of gyroscopic rigidity (similar to that provided in the spin-stabilized spacecraft 20 of FIG. 1 by the spinning spacecraft body 22). In particular, the gyroscopic rigidity of the momentum wheel 50 limits movement about the roll and yaw axes and pitch attitude control is realized by modulating the wheel's velocity to thereby exchange momentum between the wheel and the spacecraft's body 44.
Roll attitude control is obtained with an actuator (e.g., a thruster, magnetic coil, or use of the solar wings as solar pressure sails) which can generate attitude-correction torque about this axis. In the course of the orbit, there is an interchange every 6 hours between the roll and yaw axes so that yaw control is facilitated through roll control.
In the control system of FIG. 3A, the amplitude of the angular momentum H can be altered but not its orientation. This single degree of freedom is expanded to two degrees of freedom in FIG. 3B by adding a second momentum wheel 51 and inclining the wheel's axes relative to each other. The wheels 50 and 51 now have momentums H.sub.1 and H.sub.2 respectively. This system permits the orientation of the combined angular momentum to be changed so that the attitude of a spacecraft element (e.g., an antenna) can be rotated to a desired inertial attitude (i.e., the control system of FIG. 3B is more agile than that of FIG. 3A).
Because the momentum wheels of FIGS. 3A and 3B maintain a significant angular momentum at all times, attitude control systems of this type are referred to as "biased-momentum systems". However, spacecraft agility is enhanced if stored momentum is kept low because this facilitates rapid attitude corrections. FIG. 3C shows a control system which includes three reaction wheels 60, 61 and 62 that are each aligned with one of three orthogonal spacecraft axes. The velocity of reaction wheels can be varied in either direction from a rest state in which they have no velocity. Accordingly, the attitude control systems of such body-stabilized spacecraft are referred to as zero-momentum systems.
With this system, disturbance torques are corrected by continually varying the wheel velocities. Because the mean wheel velocity is near zero, gyroscopic rigidity of the wheels is minimal which facilitates quick attitude changes. However, realizing this agility not only requires a reaction-wheel system but typically also requires addition of control systems with complex, costly elements (e.g., gyroscopic inertial reference units, static earth sensors, magnetometers, star sensors and control processors). In addition, a large number of thrusters are needed to unload the wheels' momentum when it becomes excessive.
Although conventional spacecraft configurations (e.g., those exemplified in FIGS. 1, 2 and 3A-3C) have successfully stabilized a large number of existing spacecraft, they have several disadvantages. Because the antennas of spinner spacecraft must also spin, they are inherently low gain structures (e.g., omni or toroidal antennas). Although gyrostat spacecraft can carry high gain antennas, both spinner and gyrostat spacecraft have inefficient solar arrays. At any given time, one half of the solar cells of their arrays are in shadow and the effectiveness of most cells of the remaining one half is reduced because they are angled away from the sun. A small portion of a gyrostat's solar cells would meet its energy needs if fully exposed to solar radiation. The weight of the additional solar cells must be subtracted from the spacecraft's payload which significantly reduces the spacecraft's revenues.
Body-stabilized spacecraft can carry high-gain antennas and efficient planar solar panels. However, momentum-biased systems lack attitude agility and zero-momentum systems require complex control systems and a large number of momentum-unloading thrusters.